Compressor

ABSTRACT

A compressor  13,  for example for a gas turbine engine ( 10,  FIG.  1 ), comprises a circumferentially extending static casing component  20  defining a generally axially extending main gas flow path  24.  The compressor also includes a rotating aerofoil arrangement  26  comprising an aerofoil support  30  and a plurality of aerofoils  28  extending radially from the aerofoil support  30  into the main gas flow path  24.  The aerofoil support  30  and the circumferentially extending static casing component  20  define a circumferentially extending generally radial passage  40  having an outlet  42  in communication with the main gas flow path  24  through which leakage air may flow into the main gas flow path  24,  and the static casing component  20  includes a continuous circumferentially extending projection formation  44  projecting into the radial passage  40  to control the flow of leakage air into the main gas flow path  24.

The present invention relates to a compressor, and in particular to a compressor for a gas turbine engine.

In axial compressors, leakage of air occurs through the seal between the upstream stationary annulus and the first rotating blade row of the first rotor stage. The leakage air, which is recirculated from a downstream compressor stage, is at a higher temperature and pressure than the air in the main gas flow path through the compressor and enters the main gas flow path with a significant radial component of velocity.

The interaction of the leakage air with the air in the main gas flow path can result in significant pressure losses at the hub of the first rotor stage. Whilst it would be desirable to completely eliminate the leakage air, this is not possible in practice and significant pressure losses at the hub of the first rotating blade row can thus occur, resulting in significant changes in compressor efficiency.

According to a first aspect of the present invention, there is provided a compressor comprising:

a circumferentially extending static casing component defining a generally axially extending main gas flow path;

a rotating aerofoil arrangement comprising an aerofoil support and a plurality of aerofoils extending radially from the aerofoil support into the main gas flow path;

the aerofoil support and the circumferentially extending static casing component defining a circumferentially extending generally radial passage having an outlet in communication with the main gas flow path through which leakage air may flow into the main gas flow path;

wherein the static casing component includes a continuous circumferentially extending projection formation projecting into the radial passage to control the flow of leakage air into the main gas flow path.

Where the terms radial, axial and circumferential are used in this specification in relation to components of the compressor, they refer to the orientation of the particular component relative to the central longitudinal axis of the compressor.

The projection formation may be located substantially at or may be located substantially adjacent to the outlet of the circumferentially extending radial passage.

The projection formation may include a first continuous circumferentially extending projection which may project into the circumferentially extending radial passage. The first continuous circumferentially extending projection may include a radially inner sloping side surface which may have a generally concave profile. The first continuous circumferentially extending projection may include a radially outer sloping side surface which may have a generally concave profile.

Either one or both of the radially inner and radially outer sloping side surfaces may have a half angle of between 5° and 30°.

The first continuous circumferentially extending projection may project across at least 10% of the width of the circumferentially extending radial passage. The first continuous circumferentially extending projection may project across between 10% and 20% of the width of the circumferentially extending radial passage.

The distance between the centreline of the first continuous circumferentially extending projection and the outlet of the circumferentially extending radial passage may be between 33% and 66% of the radial dimension (h) of a securing member of one of the plurality of aerofoils.

The projection formation may include a second continuous circumferentially extending projection which may project into the circumferentially extending radial passage.

The first and second continuous circumferentially extending projections may be located substantially adjacent to each other. The first continuous circumferentially extending projection may be located radially inwardly of the second continuous circumferentially extending projection.

The second continuous circumferentially extending projection may have a radially inner sloping side surface which may have a generally concave profile.

The second continuous circumferentially extending projection may project across between 10% and 20% of the width of the circumferentially extending radial passage.

The radially inner sloping side surface of the second continuous circumferentially extending projection may have a half angle of between 5° and 30°.

The aerofoil may comprise a rotor blade. The aerofoil support may comprise a rotor blade mounting disc.

The compressor may be a multi-stage axial compressor for a gas turbine engine.

According to a second aspect of the present invention, there is provided a gas turbine engine including a compressor according to the first aspect of the invention.

An embodiment of the present invention will now be described by way of example only and with reference to the accompanying drawings, in which:

FIG. 1 is a diagrammatic cross-sectional view of a gas turbine engine;

FIG. 2 is a diagrammatic cross-sectional view of part of a compressor according to the invention;

FIG. 3 is an enlarged view of region A of the compressor shown in FIG. 2; and

FIG. 4 is a view similar to FIG. 3 showing air flow through region A of the compressor.

Referring to FIG. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, combustion equipment 15, a high pressure turbine 16, an intermediate pressure turbine 17, a low pressure turbine 18 and an exhaust nozzle 19.

The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produces two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor 13 compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13, and the fan 12 by suitable interconnecting shafts.

FIG. 2 shows part of a compressor according to the present invention, which may be the intermediate pressure compressor 13 of the gas turbine engine 10 shown in FIG. 1. The intermediate pressure compressor 13 may be a multi-stage axial compressor and FIG. 2 shows in particular the first rotating blade row of the first stage of the compressor 13.

The compressor 13 comprises a circumferentially extending static casing component 20 in the form of an annulus 22 and this defines a generally axially extending main gas flow path 24 for directing a main flow through the compressor 13, as indicated by the arrows M. The compressor 13 also comprises a rotating aerofoil arrangement 26 comprising a plurality of aerofoils 28 (one of which is shown in FIG. 2) which are circumferentially spaced about the central longitudinal axis X-X of the compressor 13 (equivalent to the engine axis X-X of FIG. 1) and which extend into the main gas flow path 24 radially from an aerofoil support 30 in the form of a blade rotor disc 32.

The aerofoil support 30 includes a rotating sealing arrangement 34 in the form of a knife seal which contacts a corresponding static seal member 36 on the static casing component 20. As is clear from FIG. 2, the static casing component 20, and in particular a generally radially extending part 38 thereof, and the aerofoil support 30 together define a circumferentially extending substantially radial passage 40 having an outlet 42 in communication with the main gas flow path 24 through the compressor 13.

In use, air from a subsequent compressor stage is recirculated through the compressor 13 to maintain effective sealing contact between the rotating sealing arrangement 34 and the static seal member 36. As this recirculated air is at a higher pressure and temperature than the air in the main gas flow path 24 and as the seal between the rotating sealing arrangement 34 and the static seal member 36 is not a perfect seal, leakage air tends to flow through the radial passage 40 and into the main gas flow path 24 via the outlet 42, as indicated by the arrows L.

In accordance with embodiments of the invention, the static casing component 20, and in particular the radially extending part 38 thereof, include a continuous circumferentially extending projection formation 44 which projects into the circumferentially extending radial passage 40, desirably substantially at or substantially adjacent to the outlet 42 thereof. This controls the flow of leakage air from the radial passage 40 into the main gas flow path 24, as will be described in further detail later in the specification.

In more detail and referring in particular to FIGS. 2 and 3, the projection formation 44 is located substantially at the outlet 42 of the radial passage 40 into the main gas flow path 24. In embodiments of the invention, the projection formation 42 comprises first and second projections 46, 48, with the first projection 46 being located radially inwardly of the second projection 48.

The first projection 46 comprises radially inner and radially outer sloping side surfaces 46 a, 46 b and a generally planar upper surface 46c. Each of the sloping side surfaces 46 a, 46 b has a generally concave profile and has a half angle (ε₁) of between 5° and 30°. The distance (d₁) by which the first projection 46 extends across the width (l) of the radial passage 40 is at least 10% of the width (l) of the radial passage 40, and may be between 10% and 20% of the width (l) of the radial passage 40. The first projection 46 is located radially inwardly from the outlet 42 such that the distance between the outlet 42 and the centreline of the first projection 46 is between approximately 33% and 66% of the radial dimension (h) of a securing member 50 of one of the aerofoils 28.

The second projection 48 is located at the outlet 42 of the radial passage 40 and has a radially inner sloping side surface 48 a having a generally concave profile. The radially inner sloping side surface 48 a also has a half angle (ε₂) of between 5° and 30°, and in embodiments of the invention, the half angle (ε₂) of the radially inner sloping side surface 48 a of the second projection 48 is equal to the half angle (ε₁) of the radially inner sloping side surface 46 a of the first projection 46. The distance (d₂) by which the second projection 48 extends across the width (l) of the radial passage 40 is at least 10% of the width (l) of the radial passage 40, and may be between 10% and 20% of the width (l) of the radial passage 40. In embodiments of the invention, the distance (d₂) is equal to the distance (d₁).

Referring to FIGS. 2 and 4, when the compressor 13 according to the invention is in operation, as indicated above air flows along the generally axial main gas flow path 24 as indicated by the arrows M and leakage air flows along the radial passage 40, towards the outlet 42 and into the main gas flow path 24, as indicated by the arrows L. Due to the presence of the projection formation 44, and in particular the first and second adjacent projections 46, 48, a low pressure region LP is created in the radial passage 40, towards one side thereof, generally between the first projection 46 and the outlet 42. The pressure in this low pressure region LP is lower than the pressure in the main gas flow path 24, and this creates a suction or recirculation zone into which air flowing along the main gas flow path 24 is drawn. The suction or recirculation zone created by this low pressure region LP reduces the effective area of the outlet 42 into the main gas flow path 24, causing deviation of the leakage flow towards the opposite side of the radial passage 40 at which the outlet 42 is not obstructed and causing the formation of a wake in region W by the leakage flow.

The presence of the wake is believed to reduce the rate of the leakage flow into the main gas flow path 24, thereby decreasing the radial component of velocity of the leakage flow and the radial angle of the leakage flow. By reducing the component of radial velocity and the radial angle of the leakage flow into the main gas flow path 24, disturbance to the air flowing through the main gas flow path 24 is minimised, thereby reducing the effect on the operating efficiency of the compressor 13.

The geometry of the first and second projections 46, 48 is believed to be of particular importance in being able to adequately control the flow of leakage air from the radial passage 40 through the outlet 42 into the main gas flow path 24. In particular, it is important that the radially inner sloping side surface 46 a of the first projection 46 has a concave profile as this tends to cause deviation of the flow of leakage air towards the unobstructed side of the outlet 42 of the radial passage 40, as illustrated diagrammatically in FIG. 4.

There is thus provided a compressor in which the flow of leakage air into the main gas flow path can be controlled to minimise disturbance to air flowing through the main gas flow path, thereby minimising the risk of separation of the boundary layer from the aerofoil and resultant changes in compressor efficiency.

Although embodiments of the invention have been described in the preceding paragraphs with reference to various examples, it should be appreciated that various modifications to the examples given may be made without departing from the scope of the present invention, as claimed. For example, the projection formation 44 may comprise only one of the first or second projections 46, 48. The position and dimensions of the first and/or second projections 46, 48 may vary dependent upon the particular application of the compressor and/or its dimensions and/or its operating regime. 

1. A compressor comprising: a circumferentially extending static casing component defining a generally axially extending main gas flow path; a rotating aerofoil arrangement comprising an aerofoil support and a plurality of aerofoils extending radially from the aerofoil support into the main gas flow path; the aerofoil support and the circumferentially extending static casing component defining a circumferentially extending generally radial passage having an outlet in communication with the main gas flow path through which leakage air may flow into the main gas flow path; wherein the static casing component includes a continuous circumferentially extending projection formation projecting into the radial passage to control the flow of leakage air into the main gas flow path.
 2. A compressor according to claim 1, wherein the projection formation is located substantially at or substantially adjacent to the outlet of the circumferentially extending radial passage.
 3. A compressor according to claim 1, wherein the projection formation includes a first continuous circumferentially extending projection projecting into the circumferentially extending radial passage.
 4. A compressor according to claim 3, wherein the first continuous circumferentially extending projection includes a radially inner sloping side surface having a generally concave profile.
 5. A compressor according to claim 3, wherein the first continuous circumferentially extending projection includes a radially outer sloping side surface having a generally concave profile.
 6. A compressor according to claim 4, wherein either one or both of the radially inner and radially outer sloping side surfaces have a half angle of between 5° and 30°.
 7. A compressor according to claim 3, wherein the first continuous circumferentially extending projection projects across at least 10% of the width of the circumferentially extending radial passage.
 8. A compressor according to claim 3, wherein the first continuous circumferentially extending projection projects across between 10% and 20% of the width of the circumferentially extending radial passage.
 9. A compressor according to claim 3, wherein the distance between the centreline of the first continuous circumferentially extending projection and the outlet of the circumferentially extending radial passage is between 33% and 66% of the radial dimension of a securing member of one of the plurality of aerofoils.
 10. A compressor according to claim 3, wherein the projection formation includes a second continuous circumferentially extending projection projecting into the circumferentially extending radial passage.
 11. A compressor according to claim 10, wherein the first and second continuous circumferentially extending projections are located substantially adjacent to each other.
 12. A compressor according to claim 10, wherein the first continuous circumferentially extending projection is located radially inwardly of the second continuous circumferentially extending projection.
 13. A compressor according to claim 10, wherein the second continuous circumferentially extending projection has a radially inner sloping side surface having a generally concave profile.
 14. A compressor according to claim 10, wherein the second continuous circumferentially extending projection projects across between 10% and 20% of the width of the circumferentially extending radial passage.
 15. A compressor according to claim 10, wherein the radially inner sloping side surface of the second continuous circumferentially extending projection has a half angle of between 5° and 30°.
 16. A compressor according to claim 1, wherein aerofoil comprises a rotor blade.
 17. A compressor according to claim 1, wherein the aerofoil support comprises a rotor blade mounting disc.
 18. A compressor according to claim 1, wherein the compressor is a multi-stage axial compressor for a gas turbine engine.
 19. A gas turbine engine including a compressor according to claim
 1. 